The present invention generally relates to gas turbine engines, and in particular to a device and method for cooling an engine combustor liner.
Gas turbine engines typically include a compressor for supplying pressurized air to a combustion zone in which the pressurized air is mixed with fuel and burned to generate hot combustion gases for powering the turbine. A combustor liner, which may be formed from a metallic sheet, is typically provided to protect surrounding engine structure from the high-temperature combustion gases. Combustor liners are cooled to increase the life of the liner.
Some cooling of the combustor liner may be accomplished by directing a portion of the compressor air to flow over an exterior, or “cold,” surface of the combustor liner and remove heat by convection. A conventional combustor liner may include an array of very small openings, referred to as effusion holes, which allow a portion of the cooling air to enter the combustor liner and to also remove heat by convection.
A conventional turbine engine assembly 10, partially shown in the cross-sectional view of FIG. 1, may include a compressor 11, a combustor 13, and a fuel nozzle 15 extending into a combustion zone 17 located within the combustor 13. The combustion zone 17 may be partially enclosed by a combustor liner 20 which functions to reduce the amount of heat emanating from the combustion zone 17 onto casing surfaces of the engine assembly 10. The combustor liner 20 may comprise one or more non-planar sheets of metal or other heat-resistant material. In the configuration shown, the combustor liner 20, which may be generally toroidal in shape, comprises a contoured (i.e., nonplanar) outer liner 21, a contoured inner liner 23, and a liner dome 25. It should be understood that the outer liner 21 and the inner liner 23 can be, for example, cylindrical or conical in shape.
The compressor 11 may supply a cooling air stream 31 which is incident upon the combustor liner 20. A first portion of the cooling air stream 31 may pass through effusion holes (not shown) in the liner dome 25 and enter the combustion zone 17 as a liner dome air stream 33. Additionally, a second portion of the cooling air stream 31 may pass through effusion holes (not shown) in the outer liner 21 and enter the combustion zone 17 as an outer liner air stream 35. Similarly, a third portion of the cooling air stream 31 may pass through effusion holes (not shown) in the inner liner 23 and enter the combustion zone 17 as an inner liner air stream 37.
The liner dome air stream 33, the outer liner air stream 35, and the inner liner air stream 37 may be utilized in the combustion of fuel in the combustion zone 17 and subsequently flow out of the combustion zone 17 as an exhaust gas stream 39. Additional air may be supplied to the combustion zone 17 through a plurality of primary holes and dilution holes (not shown for clarity of illustration), as known in the art, on both the outer liner and the inner liner 23. This additional air is primarily utilized in the combustion process and is not considered part of the process for cooling the combustor liner 20. It can be appreciated by one skilled in the art that the combustor liner 20 may be shaped so as to aid in producing an effective mixing of fuel and air within the combustion zone 17 for efficient combustion.
Referring to FIG. 2, also of the prior art, a portion of the liner dome 25 may have an array of downstream-pointed effusion holes 41 therethrough. The liner dome 25 may include an exterior cold dome surface 27 which is positioned in the cooling air stream 31 and an interior hot dome surface 29 which partially encloses the ongoing exothermic reaction in the combustion zone 17.
In the particular configuration shown, the plurality of downstream-pointed effusion holes 41 may form a series of adjacent rows, such as rows 43a–e, each downstream-pointed effusion hole 41 having a longitudinal axis 47 which may be generally aligned with the direction of the main flow of gases in the combustion zone 17, the direction indicated by arrow 19, proximate the hot dome surface 29. That is, the particular geometry of the downstream-pointed effusion holes 41, as well as the location and number of rows 43a–e, may be determined from thermodynamic aspects of the dynamic flow of air and fuel in the combustion zone 17. As the liner dome air stream 33 (not shown in FIG. 2 for clarity) passes through the array of downstream-pointed effusion holes 41, cooling air flows along the hot dome surface 29.
FIG. 3 is an enlarged detail view of the hot surface 29, also according to the prior art, showing four of the downstream-pointed effusion holes 41 located in the adjacent rows 43a and 43b of the liner dome 25. Each longitudinal axis 47 may be oriented generally with the direction of the main flow, indicated by arrow 19. As the cooling air stream 31 impinges on the cold dome surface 27 of the liner dome 25, a portion of the cooling air stream 31 may pass through the plurality of downstream-pointed effusion holes 41 in the direction of the respective longitudinal axes 47 and enter the combustion zone 17 as the liner dome air stream 33. After passing through the downstream-pointed effusion holes 47, the liner dome air stream 33 mixes with and becomes part of the main flow. This is a consequence of the orientation of the downstream-pointed effusion holes 41 with the direction of the main flow in the combustion zone 17. It can thus be appreciated that the liner dome air stream 33 may generally have such a large velocity or momentum when leaving the downstream-pointed effusion holes 41 that the fluid dynamic characteristics of the main flow are affected by the liner dome air stream 33 proximate the hot dome surface 29.
As taught in the present state of the art, effusion holes may be oriented such that the entering cooling air is directed along the “main flow,” that is, along the local prevailing flow of hot gases in the combustion zone. For example, U.S. Pat. No. 5,129,231 issued to Becker et al. discloses a heat shield for fuel nozzles mounted at the dome of an annular combustor for a gas turbine engine, in which the effusion holes are oriented to inject the cooling air so as to be compatible with the direction of swirling air in the combustion zone.
U.S. Pat. No. 5,918,467 issued to Kwan discloses a heat shield for a gas turbine annular combustion zone in which an array of effusion holes in the heat shield is subdivided into sectors. Within each sector, the effusion holes are arranged parallel to one another and extend in the direction of the enclosed fuel consumption air swirl. U.S. Pat. No. 6,408,629 issued to Harris et al. discloses a multi-hole combustor liner in which the orientation of a select group of effusion holes is generally in the direction of the main flow, but may be altered to direct cooling air to “hot spot” regions, such as regions downstream of dilution holes. The problem remains that, because the orientation of the effusion holes is in the direction of the main flow, as taught both in Kwan '467 and in Harris et al. '629, the cooling air is introduced into the combustion zone at a speed sufficient to affect the main flow, the cooling air entrains hot gases from the combustion zone and is thus less effective at removing heat from the hot surface of the combustor liner, resulting in higher combustor liner temperatures.
As can be seen, there continues to be a need for a method and apparatus for controlling the cooling air flowing into a combustor liner.